TECHNICAL FIELD
[0001] The following disclosure relates generally to gas turbine engines and, more particularly,
to turbine nozzles and other gas turbine engine components having stress relief slots
filled with high temperature sealing material, as well as to methods for fabricating
gas turbine engine components having sealed stress relief slots.
BACKGROUND
[0002] Gas turbine engines are commonly produced to include turbine nozzles, which accelerate
and turn combustive gas flow toward the blades of a turbine rotor downstream of the
nozzle. The turbine nozzle may have a generally annular or ring-shaped body including
an inner endwall, an outer endwall circumscribing the inner endwall, and a series
of circumferentially-spaced vanes extending between the inner and outer endwalls.
The inner endwall, the outer endwall, and the vanes define a number of combustive
gas flow paths through the turbine nozzle, which conduct hot combustive gas flow during
operation of the gas turbine engine. While portions of the nozzle are exposed to combustive
gas flow during engine operation, other portions of the turbine nozzle body and its
associated mounting features are bathed in relatively cool airflow bled from a cold
section of the engine and directed along an outer cooling flow path. In certain cases,
undesired leakage can occur across the turbine nozzle interface between the outer
cooling flow path and the core gas flow path. Such leakage can negatively affect the
efficiency of the gas turbine engine, especially when smaller in size, and may increase
the volume of airflow required for cooling purposes.
[0003] Leakage across the turbine nozzle mounting interfaces can be reduced through the
usage of annular compression seals, such as flexible, pressure-activated metal seals.
Such seals may be compressed between the mounting features of the turbine nozzle (e.g.,
rails extending radially from the opposing ends of the nozzle) and neighboring static
structures within the engine. Temperature limitations may require that such compression
seals are radially offset from the core gas flow path by a certain distance to reduce
the operational temperatures to which the seals are exposed. The turbine nozzle rails
may thus be elongated in a radial direction to allow such a radial offset between
the compression seals and the core gas flow path. Unfortunately, this also has the
effect of increasing temperature differentials that develop across the radially-elongated
rails during engine operation, which may result in excessively high hoop stresses
within the rails thereby hastening Thermomechanical Fatigue (TMF) and reducing the
service lifespan of the turbine nozzle. TMF within the turbine nozzle rails may be
alleviated through the formation of stress relief slots at strategic locations in
the nozzle rail. The inclusion of stress relief slots in the nozzle rail may, however,
permit an undesirably large amount of leakage across the turbine nozzle mounting interfaces
thereby defeating the purpose of the compression seals or at least diminishing the
effectiveness thereof.
[0004] It is thus desirable to provide embodiments of a turbine nozzle having stress relief
slots formed at one or more circumferential locations in the radially-elongated rails
or similar mounting features, which reduce TMF within the turbine nozzle while also
minimizing leakage across the turbine nozzle mounting interfaces. More generally,
it would be desirable to produce embodiments of a gas turbine engine component, such
as a turbine nozzle or a combustor liner, including stress relief slots providing
the above-noted benefits. Finally, it would be desirable to provide embodiments of
a gas turbine engine employing such a gas turbine engine component, as well as methods
for fabricating such a gas turbine engine component. Other desirable features and
characteristics of the present invention will become apparent from the subsequent
Detailed Description and the appended Claims, taken in conjunction with the accompanying
Drawings and the foregoing Background.
US2007/166154 describes a gas turbine having a nozzle with stress relief slots.
BRIEF SUMMARY
[0005] Embodiments of a gas turbine engine are provided according to the appended claims
1 to 8, and 10.
[0006] Still further provided are embodiments of a method for fabricating a gas turbine
according to the appended claim 9.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] At least one example of the present invention will hereinafter be described in conjunction
with the following figures, wherein like numerals denote like elements, and:
FIG. 1 is a schematic of an exemplary gas turbine engine including one or more turbine
nozzles;
FIG. 2 is an isometric cutaway view of a turbine nozzle (partially shown) suitable
for usage within the gas turbine engine shown in FIG. 1, which has a plurality of
sealed stress relief slots formed therein and which is illustrated in accordance with
an exemplary embodiment of the present invention;
FIG. 3 is a cross-sectional view of the turbine nozzle shown in FIG. 2 illustrating
one manner in which the turbine nozzle may be positioned between high and low pressure
turbine stages when installed within a gas turbine engine; and
FIGs. 4 and 5 are front views of a sealed stress relief slot included within the exemplary
turbine nozzle shown in FIGs. 2 and 3, as illustrated after and prior to filling with
a high temperature sealing material, respectively.
[0008] For simplicity and clarity of illustration, the drawing figures illustrate the general
manner of construction, and descriptions and details of well-known features and techniques
may be omitted to avoid unnecessarily obscuring the invention. Additionally, elements
in the drawings figures are not necessarily drawn to scale. For example, the dimensions
of some of the elements or regions in the figures may be exaggerated relative to other
elements or regions to help improve understanding of embodiments of the invention.
DETAILED DESCRIPTION
[0009] The following Detailed Description is merely exemplary in nature and is not intended
to limit the invention or the application and uses of the invention. Furthermore,
there is no intention to be bound by any theory presented in the preceding Background
or the following Detailed Description. Terms such as "comprise," "include," "have,"
and variations thereof are utilized herein to denote non-exclusive inclusions. Such
terms may thus be utilized in describing processes, articles, apparatuses, and the
like that include one or more named steps or elements, but may further include additional
unnamed steps or elements.
[0010] FIG. 1 is a simplified cross-sectional view of a gas turbine engine (GTE)
20 illustrated in accordance with an exemplary embodiment of the present invention.
By way example, GTE
20 is illustrated in FIG. 1 as a two spool turbofan engine including an intake section
22, a compressor section
24, a combustion section
26, a turbine section
28, and an exhaust section
30. Intake section
22 includes an intake fan
32 mounted in a nacelle assembly
34. In the illustrated example, compressor section
24 includes a single compressor
36, which is rotatably disposed within an engine case
38 mounted within nacelle assembly
34. Turbine section
28 includes a high pressure (HP) turbine rotor
40 and a low pressure (LP) turbine rotor
42, which are rotatably disposed within engine case
38 in flow series. An HP turbine nozzle
43 is disposed immediately upstream of HP turbine rotor
40, and an LP turbine nozzle
45 is likewise disposed upstream of LP turbine rotor
42. Compressor
36 and HP turbine rotor
40 are mounted to opposing ends of an HP shaft
44, and intake fan
32 and LP turbine rotor
42 are mounted to opposing ends of a LP shaft
46. LP shaft
46 and HP shaft
44 are co-axial; that is, LP shaft
46 extends through a longitudinal channel provided through HP shaft
44. Engine case
38 and nacelle assembly
34 terminate in a mixer nozzle
48 and a propulsion nozzle
50, respectively. Mixer nozzle
48 cooperates with a centerbody
52 to form an exhaust mixer
54, which mixes hot combustive gas flow received from turbine section
28 with cooler bypass airflow during operation of GTE
20.
[0011] As illustrated in FIG. 1 and described herein, GTE
20 is provided by way of example only. It will be readily appreciated that turbine rotors
or other metallurgically-consolidated turbine engine components of the type described
herein can be utilized within various other types of gas turbine engine including,
but not limited to, other types of turbofan, turboprop, turboshaft, and turbojet engines,
whether deployed onboard an aircraft, watercraft, or ground vehicle (e.g., a tank),
included within an auxiliary power unit, included within industrial power generators,
or utilized within another platform or application. With respect to exemplary GTE
20, in particular, it is noted that the particular structure of GTE
20 will inevitably vary amongst different embodiments. For example, in certain embodiments,
GTE
20 may include an exposed intake fan (referred to as an "open rotor configuration")
or may not include an intake fan. In other embodiments, GTE
20 may employ centrifugal compressors or impellers in addition to or in lieu of axial
compressors. In still further embodiments, GTE
20 may include a single shaft or three or more shafts along with varying numbers of
compressors and turbines.
[0012] During operation of GTE
20, air is drawn into intake section
22 and accelerated by intake fan
32. A portion of the accelerated air is directed through a bypass flow passage
56, which is provided between nacelle assembly
34 and engine case
38 and conducts relatively cool airflow over and around engine case
38. The remaining portion of air exhausted from intake fan
32 is directed into compressor section
36 and compressed by compressor
36 to raise the temperature and pressure of the core airflow. The hot, compressed airflow
is supplied to combustion section
26 wherein the air is mixed with fuel and combusted utilizing one or more combustors
58 included within section
26. The combustive gasses expand rapidly and flow through turbine section
28 to rotate the turbine rotors of HP turbine rotor
40 and LP turbine rotor
42. HP turbine nozzle
43 further accelerates the combustive gas flow and helps to impart the gas flow with
a desired tangential component prior to reaching HP turbine rotor
40. Similarly, LP turbine nozzle
45 receives the gas flow discharged from HP turbine rotor
40, accelerates and turns the gas flow toward the blades of LP turbine rotor
42. The rotation of turbine rotors
40 and
42 drives the rotation of shafts
44 and
46, respectively, which, in turn, drives the rotation of compressor
36 and intake fan
32. The rotation of shafts
44 and
46 also provides significant power output, which may be utilized in a variety of different
manners, depending upon whether GTE
20 assumes the form of a turbofan, turboprop, turboshaft, turbojet engine, or an auxiliary
power unit, to list but a few examples. After flowing through turbine section
28, the combustive gas flow is then directed into exhaust section
30 wherein mixer
54 mixes the combustive gas flow with the cooler bypass air received from bypass flow
passages
56. Finally, the combustive gas flow is exhausted from GTE
20 through propulsion nozzle
50.
[0013] FIG. 2 is an isometric cutaway view of a turbine nozzle
60 (partially shown), as illustrated in accordance with an exemplary embodiment of the
present invention. Turbine nozzle
60 can be utilized as HP turbine nozzle
43 or as LP turbine nozzle
45 shown in FIG. 1. Turbine nozzle
60 includes an annular or ring-shaped body comprised of an outer ring or endwall
62, an inner ring or endwall
64, and a plurality of airfoils or vanes
66. While only a limited portion of nozzle
60 is shown in FIG. 2, it will be appreciated that endwalls
62 and
64 are annular structures, which are generally axisymmetric with respect to the centerline
of nozzle
60 and which extend fully therearound (and, thus, around the rotational axis of GTE
20 when nozzle
60 is installed therein). Nozzle vanes
66 extend radially between outer endwall
62 and inner endwall
64 to define a number of combustive gas flow paths
68 through the body of turbine nozzle. Each gas flow path
68 is defined by a different pair of adjacent or neighboring vanes
66; an inner surface of outer endwall
62 located between the neighboring vanes
66, as taken in a radial direction; and an interior surface region of inner endwall
64 located between the neighboring vanes
66, as taken in a radial direction. Gas flow paths
68 extend through turbine nozzle
60 in axial and tangential directions to guide combustive gas flow through the body
of nozzle
60, while turning the gas flow toward the blades of a turbine rotor downstream thereof.
Gas flow paths
68 may constrict or decrease in cross-sectional flow area when moving in a fore-aft
direction along which combustive gas flows during engine operation. Each flow path
68 thus serves as a convergent nozzle to meter and accelerate combustive gas flow through
turbine nozzle
60.
[0014] Turbine nozzle
60 is fabricated to further include mounting features facilitating installation of nozzle
60 within a gas turbine engine. For example, as indicated in FIG. 2, turbine nozzle
60 may be fabricated to include a leading or forward rail
70 and a trailing or aft rail
72. Forward rail
70 projects radially outward from a forward edge portion of outer endwall
62, while trailing or aft rail
72 projects radially outward from the opposing trailing edge portion of endwall
62. Nozzle rails
70 and
72 are generically referred to herein as "radially-extending walls," as are any structures
that project radially outwardly from the body of a gas turbine engine component. Rails
70 and
72 are advantageously formed as annular structures extending entirely around the forward
and aft edges of outer endwall
62, respectively. In the illustrated embodiment, rail
70, rail
72, and outer endwall
62 are formed as a single piece or monolithic structure, which extends around the centerline
of nozzle
60 to form an unbroken or continuous 360° hoop. However, in further embodiments, such
as when turbine nozzle
60 is produced as a segmented turbine nozzle (described below), rail
70, rail
72, and outer endwall
62 can be comprised of a number of arc-shaped pieces, which are assembled to form a
segmented annular structure extending around the centerline of nozzle
60. In this case, feather seals or other seals can be disposed between the mating interfaces
of the arc-shaped pieces to help minimize leakage across turbine nozzle
60.
[0015] Nozzle rails
70 and
72 may be integrally formed with outer endwall
62 as, for example, as a single cast piece. More generally, turbine nozzle
60 may itself be produced as a single cast and machined piece or, perhaps, produced
utilizing multiple cast pieces. In this latter regard, turbine nozzle
60 may be fabricated as a brazed turbine nozzle wherein endwall
62, endwall
64, and vanes
66 are cast as separate pieces, which are subsequently assembled and bonded to yield
the finished nozzle
60. In further embodiments, turbine nozzle
60 can be produced as a bi-cast turbine nozzle wherein vanes
66 are first cast, arranged in their desired positions, and endwalls
62 and
64 are then cast thereover using an investment casting process. In further embodiments,
multiple wedge-shaped or arc-shaped pieces are cast and subsequently bolted together
or otherwise assembled to produce the completed turbine nozzle (commonly referred
to as a "segmented turbine nozzle"). Each arc-shaped piece may include a segment of
the outer endwall, a segment of the inner endwall, and a number of vanes (typically
two to three vanes) extending therebetween. Thus, when assembled, the arc-shaped pieces
collectively form an annular turbine nozzle similar to that shown in FIG. 2, but with
mating interfaces between neighboring sections of the turbine nozzle. In this case,
nozzle rails
70 and
72 may comprise multiple sections, which may or may not contact. The foregoing examples
notwithstanding, various other fabrication techniques can also be utilized to produce
turbine nozzle
60.
[0016] FIG. 3 is a cross-sectional view of turbine nozzle
60 illustrating one manner in which nozzle
60 may be mounted within a gas turbine engine, such as GTE
20 shown in FIG. 1. In this particular example, nozzle
60 is disposed between an upstream turbine stage
76 and a downstream turbine stage
78. Upstream turbine stage
76 may include a turbine rotor having a number of blades
80 (one of which is partially shown in FIG. 3), which are circumscribed or surrounded
by a first turbine shroud
82. Similarly, downstream turbine stage
78 may likewise include a turbine rotor having a number of blades
84 (again, one of which is partially shown) circumscribed by a second turbine shroud
86. Turbine shrouds
82 and
86 are static components, which are bolted or otherwise affixed to static mounting features
included within the engine infrastructure. Two such static mounting features
88 and
90 are shown in FIG. 3 and engaged by turbine shrouds
82 and
86, respectively. As indicated in FIG. 3 by arrows
92, a core gas flow path extends through turbine stage
76, turbine nozzle
60, and turbine stage
78. Collectively, turbine shroud
82, turbine shroud
86, and turbine nozzle
60 partition or separate the core gas flow path from a secondary cooling flow path,
which is located radially outboard of the core gas flow path. As further indicated
by arrows
94, the secondary cooling flow path conducts relatively cool airflow bled from an upstream
cold section of the engine and supplied to components within the hot section of the
engine for cooling purposes.
[0017] Leakage between the secondary cooling flow path and the core gas flow path may occur
at the interfaces between the turbine nozzle
60, mounting feature
88, and mounting feature
90 if not adequately sealed. In the case of larger gas turbine engines, such leakage
may have relatively little impact on engine performance. However, in the case of smaller
gas turbine engine platforms, leakage between the secondary cooling and core gas flow
paths can have an appreciable impact on overall engine performance. Additionally,
leakage between the secondary cooling and core gas flow paths can increase the volume
of airflow bled from the cold section and directed along secondary cooling flow path
94 for cooling purposes. Annular compression seals can be utilized to significantly
reduce such leakage. For example, as shown in FIG. 3, a first annular compression
seal
96 may be positioned between static mounting feature
88 and forward rail
70 of turbine nozzle
60, while a second annular compression seal
98 may be positioned between static mounting feature
90 and aft rail
72 of nozzle
60. As generally illustrated in FIG. 3, annular compression seals
96 and
98 may be pressure-activated metal seals having convolute cross-sectional geometries.
Compression seals of this type are highly effective at minimizing or eliminating leakage
across the turbine nozzle mounting interfaces. In further embodiments, seals
96 and
98 may assume other forms suitable for forming annular gas-to-gas seals between the
turbine nozzle rails and their associated mounting features.
[0018] While effective at impeding gas flow leakage, annular compression seals
96 and
98 may be associated with temperature limitations requiring compression seal
96 and/or seal
98 to be radially offset from the core gas flow path
92. For example, and with continued reference to the exemplary embodiment shown in FIGs.
2 and 3, the operational temperatures to which seal
96 is subjected if disposed in close proximity to the core gas flow path
92 (indicated in FIG. 3 in phantom) may be undesirably high. Consequently, as indicated
by arrow
99, compression seal
96 may be moved (by design) to a more remote position radially offset from gas path
92, which is heated to somewhat lower temperatures during engine operation. As further
indicated in FIG. 3, the radial length or height of forward nozzle rail
70 (identified as "H
FR") is increased to allow compression seal
96 to be moved radially outward in this manner. However, in further embodiments, the
radial height of aft rail
72 (identified as "H
AR") may be increased in essentially the same manner as is the height of forward rail
70 to allow a radial offset between compression seal
98 and gas flow path
92.
[0019] As the radial height (H
FR) of forward nozzle rail
70 increases, so too does the temperature differential that develops across rail
70 during engine operation. Undesirably rapid TMF may consequently occur within nozzle
rail
70 and the neighboring regions of turbine nozzle
60 if the resultant thermomechanical stress is not addressed. For this reason, a plurality
of stress relief slots
74 may be formed through an outer annular region of nozzle rail
70. Stress relief slots
74 may be angularly spaced about the centerline of nozzle
60 at substantially regular intervals; however, this need not always be the case. FIG.
4 illustrates one stress relief slot
74 in greater detail. Referring collectively to FIGs. 2-4, stress relief slots
74 extend axially through an outer annular portion
100 of forward nozzle rail
70 (identified in FIG. 4) and terminate adjacent inner annular portion
102 of rail
70 (also identified in FIG. 4). Outer annular portion
100 of forward rail
70 remains relatively cool during engine operation, while inner annular portion
102 of rail
70 is heated to relatively high temperatures. Absent stress relief slots
74, the outer radial growth of inner annular portion
102 is restricted by outer annular portion
100 and relatively rapid TMF may result. Stress relief slots
74 allow the outer annular region of rail
70 to better accommodate the outward radial growth of inner annular portion
102 (essentially by breaking the tensile hoop stress within outer annular portion
100) thereby reducing compressive hoop stress within inner annular portion
102 of rail
70. Stress relief slots
74 may have any shape suitable for providing this stress relief function, such as a
J-shaped geometry (shown), an anchor-shaped geometry, or keyhole-shaped geometry.
It is generally preferred, however, that stress relief slots
74 are produced to have substantially uniform widths to facilitate filling with the
high temperature sealing material, as described more fully below.
[0020] With continued reference to FIGs. 2-4, and as shown most clearly in FIG. 3, stress
relief slots
74 extend radially inward or inboard of compression seal
96 and through the annular sealing surface of rail
70 (that is, the annular region of rail
70 contacting compression seal
96). Significant gas flow leakage may thus occur across forward rail
70 (thereby bypassing compression seal
96) if stress relief slots
74 are left open or unfilled. To minimize such leakage, stress relief slots
74 are filled or plugged with a high temperature sealing material
104 (identified in FIGs. 3 and 4 by dot stippling). Various
different types of sealing material can be utilized to plug or fill stress relief
slots
74, providing that the following criteria are met: (i) the sealing material has high
temperature properties sufficient to withstand the operating conditions within the
gas turbine engine without excessive degradation, (ii) the sealing material is able
to form a sufficiently strong bond with the interior surfaces of slots
74 to prevent dislodgement during usage of turbine nozzle
60, and (iii) the sealing material has a mechanical strength less than that of the nozzle
rail parent material to enable the sealing material to crack or fracture and relieve
thermomechanical stress in the below-described manner. Materials satisfying the aforementioned
criteria include, but are not limited to, high temperature braze materials. In one
embodiment, a nickel-based braze material containing at least one melting point depressant,
such as a relatively small weight percentage of boron, is utilized as the high temperature
sealing material.
[0021] During fabrication of turbine nozzle
60, stress relief slots
74 may be cut into forward nozzle rail
70 utilizing, for example, an Electrical Discharge Machining (EDM) wire technique. Advantageously,
such a technique may allow the respective widths of slots
74 to be minimized. For example, as indicated in FIG. 5 (which illustrates one of stress
relief slots
74 prior to filling with high temperature sealing material), slots
74 may be formed to have a width of Ws, which may be on the order of about 0.02032 to
0.0254 cm (0.008 to 0.010 inch). The chosen high temperature sealing material may
be introduced into stress relief slots
74 after slots
74 have been cut or otherwise formed in forward rail
70. According to the invention, the braze material may be disposed adjacent or within
stress relief slots
74 and then subjected to heat treatment to melt the braze, fill slots
74 with little to no voiding, and form the desired bonds between the braze material
and the interior surfaces of slots
74. In certain implementations wherein the braze material is needle-dispensed, brushed,
or otherwise applied in liquid or slurry form over stress relief slots
74, the braze material may flow into slots
74 by capillary forces prior to or during heat treatment. In other implementations,
the braze foil may be cut into flexible strips, which are then inserted into stress
relief slots
74, In this latter case, the strips of braze foil may extend beyond stress relief slots
74 to ensure a sufficient volume
of braze is present to completely fill slots
74 without voiding when subject to heat treatment. If desired, the strips of braze foil
may be augmented with braze paste.
[0022] As a temperature gradient develops across forward nozzle rail
70, hairline cracks or factures may develop within the high temperature sealing material
104 contained within stress relief slots
74. Such fractures are advantageous in the sense that they allow stress relief slots
74 to provide their primary function of alleviating thermomechanical stress within forward
rail
70 and turbine nozzle
60 during engine operation. It may be noted that a certain amount of leakage may occur
across the fractures within sealed stress relief slots
74. However, any such leakage will be a small fraction of that which would otherwise
occur if stress relief slots
74 were not filled with the high temperature sealing material. This may be more fully
appreciated by referring once again to FIG. 4, which depicts a hairline fracture
106 that may form in the body of sealing material
104 occupying the illustrated slot
74 during usage of nozzle
60. As can be seen, the width of fracture
106 (identified in FIG. 4 as "W
F") is significantly less than the overall width of slot
74 (again, identified in FIG. 5 as "W
S"). For example, the width of fracture
106 (W
F) may be less than 0.00254 cm (0.001 inch) in an embodiment and, therefore, approximately
1/8 to 1/10 the width of stress relief slot
74 (when produced to have a width of about 0.02032 to 0.0254 cm (0.008 to 0.010 inch)
as described above). Thus, even when considered in the aggregate, such fractures
106 allow relatively little leakage to flow axially across forward nozzle rail
70 and bypass compression seal
96 (FIG. 3). As a result, leakage across the turbine nozzle mounting interfaces is minimized,
and overall gas turbine engine efficiency is improved. Moreover, in at least some
cases, a minimal amount of leakage through sealed stress relief slots
74 may be beneficial by helping to purge pockets of hot combustive gas that may otherwise
remain trapped near the nozzle mounting interfaces.
[0023] The foregoing has thus provided embodiments of a gas turbine engine component including
sealed stress relief slots, which reduce thermomechanical stress while also minimizing
leakage between core gas flow and secondary cooling flow paths. In certain embodiments,
the stress relief slots may be formed in the forward and/or aft rail of a turbine
nozzle and filled with a braze material, such as a nickel-based braze material. The
high temperature sealing material is preferably selected to have a mechanical strength
less than the parent material of the nozzle rail such that the sealing material preferentially
fractures to alleviate thermomechanical stress within the rail during operation of
the gas turbine engine; the term "fracture" encompassing separations occurring along
the bond interface between the high temperature sealing material and the surfaces
of the stress relief slots. While primarily described in the context of a turbine
nozzle having one or more radially-elongated rails, it is emphasized that the sealed
stress relief slots can also be formed in other gas turbine engine component having
at least one radially-extending wall projecting from the component body and into a
secondary cooling flow path. For example, in further embodiment, the sealed stress
relief slots may be formed in a radially-extending flange provided around the aft
outlet end of a combustor liner.
[0024] While primarily described above in the context of a turbine engine component and,
specifically, a turbine nozzle. The foregoing description also provided embodiments
of a method for fabricating such a gas turbine engine component. In one embodiment,
the method includes independently fabricating, purchasing from a supplier, or otherwise
obtaining a component body having a radially-extending wall projecting therefrom.
A plurality of stress relief slots is cut into or otherwise formed in the radially-extending
wall. The plurality of stress relief slots are then filled or infiltrated with a high
temperature sealing material, which impedes leakage across the radially-extending
wall between the core gas flow path and the secondary cooling flow path. The high
temperature sealing material is selected or formulated to have a mechanical strength
less than the material from which the radially-extending wall is produced such that
the high temperature sealing material preferentially fractures to relieve thermomechanical
stress when a temperature gradient develops across the radially-extending wall.
[0025] While multiple exemplary embodiments have been presented in the foregoing Detailed
Description, it should be appreciated that a vast number of variations exist. It should
also be appreciated that the exemplary embodiment or exemplary embodiments are only
examples, and are not intended to limit the scope, applicability, or configuration
of the invention in any way. Rather, the foregoing Detailed Description will provide
those skilled in the art with a convenient road map for implementing an exemplary
embodiment of the invention. It being understood that various changes may be made
in the function and arrangement of elements described in an exemplary embodiment without
departing from the scope of the invention as set-forth in the appended Claims.